1. Field of the Invention
This invention relates to high performance tactical rocket motors and solid propellant formulations operable at high pressures with burn rates relatively insensitive to changes in pressure and propellant temperature. More particularly, this invention relates to propulsion vehicles including the high performance propellant formulations in a high strength, low inert weight casing equipped with an erosion-resistant nozzle throat.
2. Description of the Related Art
Conventional solid propellant rocket motors operate by generating large amounts of hot gases from the combustion of a solid propellant formulation stored in the motor casing. The solid propellant formulation generally comprises an oxidizing agent, a fuel, and a binder. During operation, the gases generated from the combustion of the solid propellant accumulate within the combustion chamber until enough pressure is amassed within the casing to force the gases out of the casing and through an exhaust port. The expulsion of the gases from the rocket motor and into the environment produces thrust.
Thrust is measured as the product of the total mass flow rate of the combustion products exiting the rocket multiplied by the velocity of the exiting combustion products plus the product of the change in pressure at the exit plane multiplied by the exit area.
Increasing the pressure at which the gases are expelled from the combustion chamber raises the thrust level, which in turn increases the propulsion rate of the vehicle containing the rocket motor to thereby permit the vehicle to achieve higher speeds.
Since pressure is a measurement of force per unit exit area, it follows that the gas expulsion pressure can be increased by decreasing the diameter of the rocket motor nozzle throat through which the combustion products are expelled.
Decreasing the diameter of the nozzle throat can also increase the expansion ratio of the throat. Expansion ratio is the ratio of the area of the nozzle exit located aft of the nozzle throat to the area of the nozzle throat. Conventional tactical rocket motors have expansion ratios in the range of 6 to 9. Increased expansion ratios result in higher levels of rocket performance.
With conventional solid rocket propellant formulations, as the operating pressure increases by decreasing the diameter of the nozzle throat, for example, the burn rate of the propellant also increases. The change in burn rate (R.sub.b) as a function of the pressure change is defined as the burn rate slope, n: ##EQU1## Data for determining burn rates at different pressures are typically gathered either by standard strand testing or by test motor analysis. The determination of burn rates by such testing procedures is well known in the art. Generally, conventional solid rocket propellant formulations have burn rate slopes of 0.15 ips/psi or greater.
Propellants which exhibit generally flat regions in their pressure versus burn rate curves are known as plateau propellants. Plateau propellants have generally flat regions over an operating range of at least 1,000 psi. Conventional propellants usually exhibit a dramatic positive increase in burn rate slope at pressures above about 3,000 psi, as shown in FIG. 1.
One of the problems associated with conventional propellant formulations having an exponentially increasing propellant burn rate is that an increase consumption of propellant generally increases the operating pressure, which in turn increases the risk of catastrophic failure of the rocket motor casing.
The conventional solution to avoiding catastrophic failure of the rocket motor casing is to strengthen the rocket motor casing by constructing the casings with thick walls from strong, dense materials, such as steel. This approach, however, deleteriously imparts a severe weight penalty to the vehicle. Consequently, a greater amount of thrust and an increased propellant burn rate is required to propel the vehicle at a comparable rate.
Another problem associated with the use of conventional solid propellant formulations is that the burn rate of such formulations varies in response to changes in the temperature of the propellant at ignition. Temperature sensitivity, .pi..sub.k, is a measure of the sensitivity of the motor pressure to changes in propellant bulk temperature at ignition. .pi..sub.k is defined as: ##EQU2## Motor and strand testing at various temperatures and pressures generate the data required to determine .pi..sub.k. A typical nominal ignition temperature is in the range of 70.degree. F. to 80.degree. F.; temperature sensitivity is usually measured over a range of -65.degree. F. to 160.degree. F. The effect of temperature sensitivity on rocket performance is shown in FIG. 2. Conventional propellants have temperature sensitivities in the range of 0.15%/.degree. F. or higher.
Typical rocket motors utilize nozzle throat materials that exhibit erosion during operation. These materials are selected primarily for their low cost, rather than high performance characteristics. At lower nominal operating pressure, such as those in existing tactical missiles, the rate of erosion of the nozzle throat does not result in a large performance loss. However, at operating pressures of 3000 psi and higher, use of existing nozzle throat materials results in substantially higher rates of erosion of the nozzle throat. Studies have shown that nozzle throat erosion is one of the most significant sources of performance loss, and that, not surprisingly, the magnitude of this loss increases as motor operating pressure and temperature increases. Moreover, the continuous erosion of the nozzle adds an element of unpredictability to the performance of the rocket motor.
An erosion-resistant nozzle throat material would allow high pressure motor operation at maximum performance efficiency without the expected performance limitations. Erosion-resistant materials should preferably have high melting points, and should be chemically inert to oxidizing gases or form an oxide that will reduce or inhibit further chemical erosion. Additionally, these materials must be capable of withstanding thermal shock and thermal stress and resisting extrusion. Although there have been motors developed that use non-eroding throat materials, such as tungsten, such non-eroding throats have generally been rejected in commercial use due to their relatively high expense and weight.
Most small diameter, for example, up to about 15 inches, tactical rocket motors comprise moderate to high strength steel cases. Air frame stiffness requirements of and the high operating pressures encountered during use of conventional solid propellants have driven the selection of high strength steel cases. In IM (insensitive munitions) testing, many of these steel case systems perform quite poorly, particularly when coupled with conventional HTPB/AP (hydroxy-terminated polybutadiene/ammonium perchlorate) propellants. Further, as described above, the overall weight of the solid propellant rocket motor propelled vehicle is a concern and increasing the weight of the motor case has an adverse impact on performance of the vehicle. Both lighter aluminum and titanium alloys have been investigated as possible materials for tactical motor casings above 5" diameter but have proven unsatisfactory for either effectiveness or cost reasons. There is a need for a rocket motor case optimally designed and composed of materials suitable for use with high pressure rocket motors and which fulfill the requirements for air frame stiffness, maximum motor operating pressure and IM testing.
The design and geometry of propellant grain also effect the performance characteristics of solid propellant rocket motors. Many existing tactical missile rocket motors use a boost-sustain thrust profile which starts at a high thrust level for generating large amounts of thrust necessary for lift-off or deployment, and subsequently decreases to a lower thrust to allow for a lower in-flight motor operating pressure. Thus, propellant grain designs should be capable of being tailored to achieve a thrust profile that maintains high thrust and motor pressure conditions throughout the course of flight.
It would be a significant advancement in the art to provide a solid rocket propellant formulation operable at high pressures without a high positive burn rate slope or high temperature sensitivity. A low or negative burn rate slope and low temperature sensitivity would result in propellant burn rates that are insensitive to increases in operating pressure and changes in propellant temperature and thus the propellant would operate at high pressures within a narrower, more predictable pressure range without an associated increase in propellant burn rate. Such a propellant would result in a more predictable and reliable operation of the rocket motor and vehicle.